[119] A description of
the mission analysis studies conducted for Project Mercury is given
along with specific examples for the various mission analysis
phases.

Aborted mission studies constituted about 90
percent of all mission- analysis studies conducted. These studies
were necessary from a light- safety standpoint and are considered
equally applicable to future manned spacecraft projects. It was found
that the basic mission design must be chosen in a flexible manner so
that consideration can be given to the changes in mission
constraints. Real-time computing has proved extremely valuable in
Project Mercury; however, consideration must be given to changes in
mission operational plans which cannot be effectively included in the
Real Time Computer Complex.

Introduction

The mission-analysis effort in Project Mercury
was conducted in several phases leading up to the flight missions.
These phases include the mission analysis supporting the systems
design of the spacecraft, the basic operational design of the Mercury
missions based on mission requirements and objectives, detailed
operational mission analysis for each specific flight, and the
formulation of the mission logic to be included in the computer used
for inflight real-time control of the missions.

Mission Phases

In figure 7-1 are shown the important phases
of mission-analysis studies. In the early mission-analysis phase, the
analysis was specifically for use in spacecraft system design. For
example, the maximum loads and heating conditions were determined for
structural design, and the spacecraft propulsion performance
requirements were determined leading to the design of the retrorocket
system. After the spacecraft systems were essentially designed, the
mission-analysis effort shifted to the operational phase. In this
phase the system design was reasonably fixed and the detailed mission
design was then accomplished by taking into account all of the
constraints, including spacecraft, launch-vehicle, and operational
constraints. The objective in this phase is to design a mission
within the capabilities of the actual spacecraft system developed. In
this phase of the mission some feedback into system design was made,
although these were small changes since the early design proved to be
sound.

The next mission analysis phase was in the
design of specific missions. In this case the mission analysis was
specialized to handle the aspects of a particular mission by using
the actual performance characteristics of the launch vehicle and
spacecraft being used. This phase also included the analysis for the
particular operational mission objectives and ground rules developed
for these missions.

The next phase was the real-time mission
analysis phase, which started at the beginning of the launch
countdown and lasted until the

Figure 7-1. Mission analysis sequence
diagram.

[120] vehicle was
recovered after the mission. In this, calculations were accomplished
in real time by a computer; however, the logic and equations used in
this computer were developed in the preceding operational
mission-analysis phase. Although every effort was made to anticipate
all the possibilities that could affect the flight and include them
in the real- time computer program, these possibilities were never
fully established. Therefore, mission-analysis experts were used as
flight controllers and also performed auxiliary computing using
off-line computers other than those used in the realtime computing
complex during the missions.

The next mission-analysis phase was a
postflight analysis phase in which the information obtained from
actual flights was fedback into the plans for future flights and, in
some cases resulted in system modifications to the spacecraft, the
launch vehicle, and the ground support system.

Some specific examples of mission constraints
affecting the analysis are shown in figure 7-2. Some of the
spacecraft constraints that must be considered are the performance of
the spacecraft propulsion system, the spacecraft control system
accuracies, and other system limitations. Some of the ground complex
constraints to be considered are performance (which includes the
effects of the locations of command stations and command ranges) and
system limitations. Constraints involving the launch vehicle which
had to be considered were performance, guidance accuracies, and
systems limitations. In Project

Figure 7-2. Operational mission
analysis.

Mercury the systems limitations of the launch
vehicle included heating and load restraints and the guidance radar
look angle constraint.

The operational constraints to be considered
in the are a of launch operations are range safety limits, abort
considerations, environmental considerations, landing and recovery
considerations, and human factors. Some of the environmental factors
that were considered were the effect of atmospheric and geophysic
constraints and winds. Consideration had to be given to recovery and
landing constraints for both normal and aborted missions and, in all
cases, the human tolerances to acceleration loads and motions were
considered.

Abort considerations resulted in about 90
percent of the mission- analysis studies. Studies were made to
provide flight controllers with the information as to when to
initiate aborts for maximum pilot safety. Studies were also made to
determine allowable tolerances in order to obtain safe miss distances
between the launch vehicle and the spacecraft and acceptable lateral
loads. Also of importance were the studies to determine the abort
recovery areas for all phases of the flight.

In order to illustrate some of the techniques
used and the results accomplished in the mission-analysis area, a few
specific examples from each phase will be discussed.

Design Mission Analysis

One example of the work performed in the
advanced mission analysis phase is illustrated by a study of the
immediate post-abort conditions. The selection of the escape-rocket
offset involved a compromise between high lateral loads and low miss
distances between the spacecraft and the launch vehicle in the
high-dynamic-pressure abort phase of launch. For low offset values
the probability of exceeding high lateral loads was low; however, the
probability of obtaining low miss distances was high. For high values
of the offset the opposite is true. Thus, the selection of the offset
was made on the basis of minimum combined probability of occurrence
of either events. In figure 7-3, the combined probability of
exceeding either a dangerous lateral load or an unacceptable miss
distance is shown plotted against the escape-rocket offset.

[121] Figure 7-3.
Selection of escape-rocket thrust offset.

Operational Mission
Analysis

A typical example of the operational mission
analysis was in the selection of the Mercury orbital elements. The
orbital inclination which governed the ground track for Project
Mercury was selected because the network facilities established prior
to Mercury could be used to good advantage, reentries for the first
three orbital passes and the 16th to the 18th passes occurred over
the United States, and the orbital ground track fell within the
temperate region of the world. In addition, the specific Mercury
inclination was affected by launch-abort recovery
considerations.

The orbital altitude and shape of the Mercury
orbit were selected based on launch-vehicle performance, accuracy,
and abort operational considerations. These considerations are
illustrated in figures 7-4 to 7-7. In figure 7-4 the orbital lifetime
is shown plotted against apogee altitude for given perigee altitudes.
For Project Mercury it was desired to have minimum lifetime of 36
hours for a 24-hour mission. Since the atmospheric densities at
orbital altitudes were not well-defined at the time Project Mercury
was initiated, it was believed that 21 conservative value for density
must be used for estimating lifetime. The density used in this figure
is considered to be a 3a [a = Greek letter sigma], or very
conservative, dense atmosphere. From figure 7-4, it can be noted that
for an adequate lifetime in a circular orbit at all altitude of 105
nautical miles could have been selected, or an elliptical orbit
having the same lifetime could have been selected, for example, an
orbit having an 80-mile perigee and a 170-mile apogee.

Figure 7-4. Minimum lifetimes for
elliptical orbits.

The next constraint to be considered is that
of launch-vehicle performance. In figure 7-5 the staging time is
shown plotted against the insertion or perigee altitude. The curves
shown are given for a constant orbital lifetime; that is, the apogee
altitude decreases as the insertion altitude increases. For a
constant insertion altitude the performance, or excess velocity
available above that required ([Greek letter Delta]Vmin), increases with
staging time until it reaches a peak value. For greater staging times
the performance decreases. The minimum acceptable performance curves
are shown in figure 7-5. The increment of velocity [Greek letter
Delta] V that defines the acceptable performance is the difference
between the velocity at fuel depletion and the planned velocity.
Therefore, all of the clear area in the figure would represent
acceptable orbital insertion altitudes.

The launch-vehicle guidance accuracies are
considered in figure 7-6. Since the Atlas launch vehicle used for the
Mercury program was guided by a radio guidance system, the guidance
accuracy was dependent to some extent on the radar elevation angle at
cut-off. In figure 7 6 the minimum elevation angle Emin which was
considered acceptable is shown. Again the clear area in the figure is
indicative of acceptable orbital insertion conditions. Next, however,
the operational considerations must be included. These are shown in
figure 7-7. In this [122] case the
operational consideration which affected the orbital conditions was
the requirement to avoid a landing in Africa for an abort from the
minimum acceptable velocity. In this figure the position of the line
shown is such that the spacecraft would not land in Africa if an
abort were made at the no-go velocity, with allowance for the
dispersions on the abort landing area. From figure 7-7 it may be
noted that the operational consideration significantly affects the
orbital insertion altitudes which could be used for Project
Mercury.

Figure 7-5. Effect of launch-vehicle
performance.

Figure 7-6. Effect of a launch-vehicle
guidance.

As operational experience was gained in
Project Mercury flights, confidence and knowledge in the systems made
it possible to reduce to some extent, the original guidance and
performance constraints. For example, the minimum elevation angle was
reduced after obtaining a better understanding of the effects on
guidance accuracy from operational experience with the guidance
system.

Figure 7-7. Effect of operational
constraints.

Specific Mission
Analysis

A considerable mission-analysis effort is made
in the design of each specific Mercury flight. Included in this
effort are detailed trajectory calculations for the mission,
dispersion calculations, calculations concerning aborts during all
phases of the mission, and calculations of retrograde time to be used
in the mission. When the flight day arrived, special mission-analysis
studies were performed to support the flight. These studies included
evaluating the wind effects on the loads on the launch vehicle and
determining the landing areas of the spacecraft in aborted missions
based on actual wind profiles. In figure 7-8 the effects of the
actual winds on the abort landing areas at various times of the
flight are shown for the MA-9

Figure 7-8. Effects of actual winds on
MA-9 abort landing.

[123] mission. These
calculations were made to enable the recovery forces to be positioned
prior to the launch such that they could most easily make an
emergency recovery should abort occur.

Real-Time Mission
Analysis

General Computing
Requirements

Real-time computing has proved very valuable
in Project Mercury for use in flight control and monitoring. The
basic computing requirements in real time are as follows:

(1) Powered flight.
Pertinent trajectory parameters were computed in order that the
status of the launch could be monitored for any indication of all
impending abort. The cut-off velocity was used to determine the
acceptability of the orbital parameters based on preplanned criteria.
In addition, landing points for possible aborts and radar-acquisition
data were computed.

(2) Aborted
missions. For aborted missions the
computer must be programed to select a target recovery area and if
necessary compute the time for retrofire to land within this
area.

(3) Orbit. In this phase the
orbital parameters were predicted with sufficient accuracy to
establish the minimum lifetime of the orbit, to predict the retrofire
time to land in normal and contingency recovery areas, to determine
spacecraft orbital position, to determine acquisition data for all
radar sites, and to predict the time of landing for use by recovery
forces.

(4) Reentry. During reentry
the computer program recalculates and updates the landing point and
time of landing, based on conditions at retrofire, in addition to
predicting acquisition data for reentry radar stations.

Example of Go-No-Go
Computation

The computation of the go-no-go parameters was
probably the most important of the realtime computations. The
selection of the Mercury go-no-go criteria which were used in the
real-time computing program is shown in figures 7-9 to 7-11. In
figure 7-9 the minimum energy for an acceptable Mercury mission is
illustrated. The flight-path angle at insertion is plotted against
the insertion velocity. The minimum acceptable orbit was defined as
that orbit in which the spacecraft could safely complete one orbital
pass and land. Because of the

Figure 7-9. Determination of minimum
acceptable orbit.

Figure 7-10. Determination of maximum
acceptable orbit.

Figure 7-11. Operational go-no-go
orbital-insertion criteria.

critical flight safety nature of the problem,
the minimum orbit was selected on the basis of a very conservative
drag coefficient CD and atmospheric density p [Greek letter rho]. The
symbol, (CDp)n, shown in figure 7- 12 has been normalized and
[124]
represents the ratio of the parametric drag coefficient-density
product to a nominal value of this product. Therefore, values of
(CDp)n which are greater than unity are considered to be
conservative. The 99- percent probability curve shown in figure 7-9
was the one selected for the go-no-go criteria. Therefore from a
lifetime consideration the conditions would be "go" at velocities
higher than this boundary; however, other constraints imposed a limit
at higher velocities.

Figure 7-12. Effects of actual
atmosphere on MA-9 orbital lifetime.

In figure 7-10 the determination of the
maximum energy orbit is illustrated. As the velocity is increased
above orbital velocity the apogee increases approximately 1 mile for
every 2 feet per second. When the velocity reaches a certain critical
value, an area occurs near perigee such that, if the retrorockets
were ignited, excessive heating would occur during reentry. As the
velocity increases above this value this critical area near perigee
extends over most of the orbit and another critical area for
initiation of reentry appears near apogee. At this point if reentry
were initiated, the reentry loads would become excessive. As the
velocity is further increased, a velocity is reached in which these
critical areas cover the entire orbital range and a safe reentry
would not be possible from any point in the orbit. The operational
go-no-go criteria that resulted from these constraints are shown in
figure 7-11 where the flight-path angle at cut-off is plotted against
the insertion velocity. The region for a minimum acceptable orbit
lies within the boundaries shown. For all Project Mercury missions
the cut-off velocities were well within the safe boundaries. For the
MA-9 mission, for example, the cut-off occurred within the boundary
of the symbol shown in this figure.

As was previously stated, some auxiliary
computing was performed during each mission outside of the real-time
computers. An example of this auxiliary computing is shown in figure
7-12 were the effects of the actual atmosphere on the orbital
lifetime of the MA-9 mission are shown. In figure 7-12 apogee
altitude is plotted against time. Because of the length of the MA-9
mission and the uncertainty of the density of the actual atmosphere
on the day of this flight, it was thought necessary to attempt to
determine the variation of the actual atmosphere from that used in
preflight computations. This calculation was necessary in order to
commit the mission to completing 22 passes at a predetermined time
during the flight. The lines shown in the figure are for
precalculated atmospheric densities which varied from that of the
assumed atmosphere. The symbols in this figure indicate the actual
apogee obtained during the flight and also that the actual atmosphere
was very close to that used in the preflight computations. The actual
orbital lifetime for the MA-9 mission would have been about 4.7 days
if a reentry were not initiated using the retrorockets.

Concluding Remarks

The operational experience obtained in
mission-analysis studies for Project Mercury has proved valuable for
application to other manned space-flight programs. Aborted mission
studies constituted about 90 percent of all the mission-analysis
studies conducted for Mercury. Although the results of these studies
were not required operationally, the amount of effort spent on abort
studies is necessary from a flight safety standpoint and will be
equally applicable [125] to future manned space projects. It is also evident
that the basic mission design must be chosen in a flexible and manner
so that consideration can be given to changes in the spacecraft
launch vehicle or operational constraints. Real-time computing has
proved extremely valuable in Project Mercury; however, it seems that
consideration must always be given to changes to mission operational
plans which cannot be effectively included in the real-time computing
complex. Therefore, auxiliary inflight computing probably should
always be considered.